1. Field of the Invention
The disclosed system and method relate generally to generating a relatively large, or steep, thermal gradient on a test object, and relates more particularly to a test facility for generating a steep thermal gradient across a thin wall of a test object or actual component, particularly gas turbine engine components.
2. Description of Related Art
Current thermal and mechanical fatigue load (TMF) test methods do not include steep temperature gradients across thin walls, or thin coated walls, representative engine pressures and velocities, or the ability to perform boundary layer cooling measurements. Given the extensive testing used to generate durability and life prediction models, and the cost of full-up engine tests ($6M+ each), no cost-effective way to conduct representative TMF testing is presently available. Current blade/vane design methods are also very costly, often requiring several iterations to achieve a “best effort” solution, resulting in increases of many millions of dollars and months of schedule to an engine program.
Additionally, TMF related failures in engine hot sections are a critical factor driving class A mishaps, and low total accumulated cycles (TAC) useful life, Mean Time Between Failure (MTBF), and Mean Time Between Maintenance (MTBM). Improved testing is expected to result in advances in designs for blade and vane components that are more resistant to the effects of TMF. Enhanced component designs can lead to more advanced engines that may be capable of reaching 4300 TAC life for the hot section. Improved designs may also help legacy engines extend hot section life to 5000 or perhaps 6000 TACs. Such improvements in blade and vane durability could save hundreds of millions of dollars in future maintenance cost avoidance both for military and commercial operations.
Sophisticated models have been used to determine blade stress, temperature, and life, but it has been impracticable to date to correlate these models to test data. Calibration of the model involves varying parameters independently to learn their effects. In a gas turbine engine, varying parameters may realistically be carried out over a very narrow relative range.
Current finite element analysis (FEA) software can approximate the stresses (mechanical and thermal) on blades and vanes, but without correlation of boundary conditions to empirical testing, it is difficult to know how accurate the model is and what risk the deviation from the conservative approach represents. Additionally, with a better understanding of the interplay between thermal and mechanical stresses, velocities, pressure distributions, and cooling flow effects, more advanced blade/vane models, and therefore, more advanced designs will be possible.
Traditional induction thermal mechanical testing has been used to represent an engine operational environment. Current TMF heating methods include direct resistive, fuel burner, laser and other radiant methods, electromagnetic induction and immersion heating. Elevated isothermal through-wall conditions have also been used for thermal tests on specimens formed with materials that are relevant to blade and vane components. However, these specimens are typically constructed for thermal material tests, and often have a geometry that is not necessarily representative of practical blade and vane geometry. Such tests do not wholly represent engine operating conditions and the resulting data is subject to very conservative interpretation. A major limiting factor to achieving rapid representative thermal excursions within an appropriately representative test component is the inability to attain adequate heat flux under representative conditions.
The modern gas turbine engine flow path undergoes a dramatic series of axial temperature, velocity, and pressure changes from the fan inlet to low turbine discharge. Components in the flow path withstand an extremely hostile environment. For example, typical first stage turbine components experience:                Velocities of approximately 0.8 Mach        Temperatures approaching 3,2000° F. static, and 3,0800° F. relative        Pressure of 470 psia        
Secondary cooling flow systems keep blades and vanes at lower temperatures than the main fluid path to maintain component strength and to avoid melting. Maintaining and controlling blade temperature is important to controlling transient blade growth to minimize tip rubbing and increase blade durability, and extend component life.
Attempting to control and phase thermal gradients on gas turbine engine components has so far proven highly impractical with existing methods. As a result, analytically based life prediction models have fallen well short of accurately predicting component life with the result of imposing serious financial burdens on operators. A major limiting factor to achieving rapid representative thermal excursions within the test component is the inability to attain adequate heat flux under representative conditions.